Composite material assembly, and more particularly fuselage manufacturing through the use of multi-piece sections, typically requires pre-solidification and cure of each piece prior to assemble them with splices between individual sections or portions.
The limitations of this methodology are:                a minimum two-step cure is required;        additional mechanical fasteners are required at splicing joints on primary structure components;        the methodology requires handling equipment and assembly jigs (for out-of-mold operations);        long fuselage manufacturing time;        over thickness at joints resulting in stress concentration;        increases in weight of assembly; and        surface preparation is required prior to bonding.        
Various solutions for assembly of multi-piece sections have been proposed in the prior art.
U.S. Pat. No. 7,459,048 discloses a method of manufacturing a unitary section of an aircraft fuselage including steps of disposing a thin layup mandrel element onto the outer shell surface of a cylindrical inner mandrel shell to form a mandrel with a layup surface. The method further includes steps of laying-up fibers onto the layup surface while the mandrel rotates to form a unitary pre-cured section of an aircraft fuselage.
WO 98/32589 discloses composite structures having a continuous skin formed using automated fiber placement methods. The multiple layers of fibers are placed on a fiber placement tool including a mandrel body surrounded by a bladder. Uncured composite structures are created by placing fibers around the fiber placement tool as discontinuous segments that are capable of moving or sliding in relation to each other in order to be expandable from within. The uncured structures are then expanded against the other surface of the molds by creating a vacuum between the bladder and the molds.
U.S. Pat. No. 7,325,771 discloses structures and methods for joining composite fuselage sections using spliced joints attaching a first stiffener on a first composite part as well as a second stiffener on a second composite part through a fitting. A strap is then used to splice the first and second composite parts together.
US 2006/0251847 discloses a method of joining composite elements in which the bonding is done through the thickness of fiber composite laminates in order to reduce interlaminar stresses using non-interlocking and interlocking bonds.
US 2009/0148647 discloses a method of fabricating composite structures by joining a plurality of composite modules along their edges using scarf joints instead of using advance fiber placement machines that require high capital investment and operating costs.
However, there is still a need for a system and method for fabricating composite material assemblies that facilitate assembly of parts when forming structures while minimizing assembly equipment costs.